CN111308911A - Full-function modular simulation processing system and method based on satellite attitude and orbit control - Google Patents

Full-function modular simulation processing system and method based on satellite attitude and orbit control Download PDF

Info

Publication number
CN111308911A
CN111308911A CN202010128913.7A CN202010128913A CN111308911A CN 111308911 A CN111308911 A CN 111308911A CN 202010128913 A CN202010128913 A CN 202010128913A CN 111308911 A CN111308911 A CN 111308911A
Authority
CN
China
Prior art keywords
attitude
module
matrix
battery array
satellite
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
CN202010128913.7A
Other languages
Chinese (zh)
Other versions
CN111308911B (en
Inventor
王嘉轶
陈秀梅
杜耀珂
王文妍
完备
崔佳
刘美师
陈桦
王禹
朱郁婓
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Shanghai Aerospace Control Technology Institute
Original Assignee
Shanghai Aerospace Control Technology Institute
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Shanghai Aerospace Control Technology Institute filed Critical Shanghai Aerospace Control Technology Institute
Priority to CN202010128913.7A priority Critical patent/CN111308911B/en
Publication of CN111308911A publication Critical patent/CN111308911A/en
Application granted granted Critical
Publication of CN111308911B publication Critical patent/CN111308911B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B17/00—Systems involving the use of models or simulators of said systems
    • G05B17/02—Systems involving the use of models or simulators of said systems electric

Abstract

The invention discloses a full-function modular simulation processing system and a method based on satellite attitude and orbit control, wherein the system comprises: the system comprises a dynamics module, an upper computer control module, an xPC real-time operation module and an FPGA module; the upper computer control module transmits an operation instruction to the xPC real-time operation module; the xPC real-time operation module receives the single-machine model data and the operation instruction and transmits the single-machine model data to the FPGA module according to the operation instruction; and the FPGA module obtains an inertial space attitude measurement value of the spacecraft, an output attitude angular velocity, noise and an installation matrix of a reaction flywheel according to the single-machine model data and transmits the installation matrix to the real satellite-borne single machine. The invention solves the problem of satellite full-function modular simulation design and realizes the development of semi-physical tests of the attitude and orbit control system under the condition of single alignment.

Description

Full-function modular simulation processing system and method based on satellite attitude and orbit control
Technical Field
The invention belongs to the technical field of spacecraft engineering, and particularly relates to a full-function modular simulation processing system and method based on satellite attitude and orbit control.
Background
At present, ground test verification work of a satellite attitude and orbit control system mainly depends on semi-physical test work as support. In the semi-physical test work of the satellite, the situation that the single machine cannot be completely sleeved or is not in place often occurs.
Disclosure of Invention
The technical problem solved by the invention is as follows: the system and the method overcome the defects of the prior art, provide a full-function modular simulation processing system and method based on satellite attitude and orbit control, solve the problem of satellite full-function modular simulation design, and realize the development of semi-physical tests of the attitude and orbit control system under the condition of single alignment.
The purpose of the invention is realized by the following technical scheme: a full-function modular simulation processing system based on satellite attitude and orbit control comprises: the system comprises a dynamics module, an upper computer control module, an xPC real-time operation module and an FPGA module; the upper computer control module transmits an operation instruction to the xPC real-time operation module; the xPC real-time operation module receives the single-machine model data and the operation instruction and transmits the single-machine model data to the FPGA module according to the operation instruction; and the FPGA module obtains an inertial space attitude measurement value of the spacecraft, an output attitude angular velocity, noise and an installation matrix of a reaction flywheel according to the single-machine model data and transmits the installation matrix to the real satellite-borne single machine.
In the full-function modularized simulation processing system based on satellite attitude and orbit control, single model data is obtained through a solar cell array and a satellite attitude dynamics equation during attitude control of the thruster after the antenna is unfolded.
In the above full-function modular simulation processing system based on satellite attitude and orbit control, the dynamic equation of the satellite attitude when the solar cell array and the thruster attitude control after the antenna are unfolded is as follows:
Figure BDA0002395263420000021
Figure BDA0002395263420000022
Figure BDA0002395263420000023
wherein, IsThe method comprises the following steps of (1) obtaining a rotational inertia matrix of a star in a three-axis coordinate system;
Figure BDA0002395263420000024
is the angular velocity vector of the star relative to the initial inertial coordinate system;
Figure BDA0002395263420000025
the angular velocity vectors of the left battery array and the right battery array are respectively; t issExternal moment acting on the satellite; fsls、FsrsFlexible coupling coefficients of vibration of the left battery array and the right battery array to rotation of the star body are respectively set; fals、FarsFlexible coupling coefficients of the vibration of the left battery array and the right battery array to the rotation of the battery arrays are respectively set; zetals、ζrsRespectively a left battery array modal damping coefficient and a right battery array modal damping coefficient; lambdaals、ΛarsRespectively a left battery array modal frequency matrix and a right battery array modal frequency matrix;
Figure BDA0002395263420000026
respectively a left solar cell array modal coordinate array and a right solar cell array modal coordinate array.
The above full functions based on satellite attitude and orbit controlIn the modularized simulation processing system, the attitude measurement value of the spacecraft inertial space is the pointing direction Z of an optical axis in the inertial spaceI＝CIbZbWherein, CibIs a conversion matrix of the body system to the inertial system, ZbIs represented by an optical axis under the system, Zb＝CbsZs，CbsIs a matrix of constant values, and the matrix of constant values,
Figure BDA0002395263420000027
Zsthe optical axis is expressed in the star sensor measurement coordinate system.
In the above full-function modular simulation processing system based on satellite attitude and orbit control, the output attitude angular velocity is:
Figure BDA0002395263420000028
wherein, ω isoutTo output attitude angular velocity, omegainFor inputting attitude angular velocity, omegaminTo a minimum amplitude limit, ωmaxIs the maximum amplitude limit.
In the above full-function modular analog processing system based on satellite attitude and orbit control, the formula of noise is:
distr＝A·sin(2iπ·w·t)；
where distr is noise, A is a disturbance amplitude, w is a disturbance frequency, and t is time.
In the above full-function modular simulation processing system based on satellite attitude and orbit control, the installation matrix of the reaction flywheel is obtained by the following formula:
D＝RT(RRT)-1；
wherein D is the distribution matrix and R is the mounting matrix of the reaction flywheel.
A full-function modular simulation processing method based on satellite attitude and orbit control comprises the following steps: the dynamic module transmits the single-machine model data to the xPC real-time operation module, and the upper computer control module transmits an operation instruction to the xPC real-time operation module; the xPC real-time operation module receives the single-machine model data and the operation instruction and transmits the single-machine model data to the FPGA module according to the operation instruction; the FPGA module obtains an inertial space attitude measurement value of the spacecraft, an output attitude angular velocity, noise and an installation matrix of a reaction flywheel according to the model data of the single machine, and transmits the installation matrix to the real satellite-borne single machine.
In the above full-function modular simulation processing method based on satellite attitude and orbit control, the single model data is obtained by a solar cell array and a satellite attitude dynamics equation during the attitude control of the thruster after the antenna is unfolded.
In the above full-function modular simulation processing method based on satellite attitude and orbit control, the dynamic equation of the satellite attitude when the solar cell array and the antenna are unfolded and the thruster attitude is controlled is as follows:
Figure BDA0002395263420000031
Figure BDA0002395263420000032
Figure BDA0002395263420000033
wherein, IsThe method comprises the following steps of (1) obtaining a rotational inertia matrix of a star in a three-axis coordinate system;
Figure BDA0002395263420000034
is the angular velocity vector of the star relative to the initial inertial coordinate system;
Figure BDA0002395263420000035
the angular velocity vectors of the left battery array and the right battery array are respectively; t issExternal moment acting on the satellite; fsls、FsrsFlexible coupling coefficients of vibration of the left battery array and the right battery array to rotation of the star body are respectively set; fals、FarsFlexible coupling coefficients of the vibration of the left battery array and the right battery array to the rotation of the battery arrays are respectively set; zetals、ζrsRespectively a left battery array modal damping coefficient and a right battery array modal damping coefficient; lambdaals、ΛarsRespectively a left battery array die and a right battery array dieA state frequency matrix;
Figure BDA0002395263420000036
respectively a left solar cell array modal coordinate array and a right solar cell array modal coordinate array.
In the full-function modular simulation processing method based on satellite attitude and orbit control, the attitude measurement value of the spacecraft inertial space is the pointing Z of the optical axis in the inertial spaceI＝CIbZbWherein, CibIs a conversion matrix of the body system to the inertial system, ZbIs represented by an optical axis under the system, Zb＝CbsZs，CbsIs a matrix of constant values, and the matrix of constant values,
Figure BDA0002395263420000037
Zsthe optical axis is expressed in the star sensor measurement coordinate system.
Compared with the prior art, the invention has the following beneficial effects:
the invention solves the problem of satellite full-function modular simulation design and realizes the development of semi-physical tests of the attitude and orbit control system under the condition of single alignment.
Drawings
Various other advantages and benefits will become apparent to those of ordinary skill in the art upon reading the following detailed description of the preferred embodiments. The drawings are only for purposes of illustrating the preferred embodiments and are not to be construed as limiting the invention. Also, like reference numerals are used to refer to like parts throughout the drawings. In the drawings:
fig. 1 is a block diagram of a full-function modular simulation processing system based on satellite attitude and orbit control according to an embodiment of the present invention;
fig. 2 is a functional block diagram of FPGA software of the satellite full-function modular simulation processing method according to the embodiment of the present invention;
fig. 3 is a schematic diagram of a control software interface module of the satellite full-function modular simulation processing method according to the embodiment of the present invention.
Detailed Description
Exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be embodied in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the disclosure to those skilled in the art. It should be noted that the embodiments and features of the embodiments may be combined with each other without conflict. The present invention will be described in detail below with reference to the embodiments with reference to the attached drawings.
Fig. 1 is a block diagram of a fully functional modular analog processing system based on satellite attitude and orbit control according to an embodiment of the present invention. As shown in fig. 1, the system comprises a dynamics module, an upper computer control module, an xPC real-time operation module and an FPGA module; the upper computer control module transmits an operation instruction to the xPC real-time operation module; the xPC real-time operation module receives the single-machine model data and the operation instruction and transmits the single-machine model data to the FPGA module according to the operation instruction; and the FPGA module obtains an inertial space attitude measurement value of the spacecraft, an output attitude angular velocity, noise and an installation matrix of a reaction flywheel according to the single-machine model data and transmits the installation matrix to the real satellite-borne single machine.
The input standalone model includes but is not limited to: attitude sensors such as star sensors and gyro combinations, and actuating components such as reaction flywheels.
The dynamic module takes satellite orbit/attitude information, installation parameters, view field angle values, single machine configuration parameters and data noise of the attitude sensor as single machine model data to be input to the xPC target machine real-time operation module, and the executing mechanism takes a satellite control instruction (rectangular wave enabling signal), an access state, single machine configuration parameters (nominal thrust) and data noise as single machine model data to be input to the xPC target machine real-time operation module.
The dynamics module mainly inputs in-orbit data (including in-orbit attitude and orbit data) of the satellite attitude and orbit control system to the xPC target machine, and the upper computer control module inputs an operation instruction to the xPC target machine.
For single-machine model data of the dynamics module, before the solar cell array is expanded, the dynamic equation of the satellite attitude is
Figure BDA0002395263420000051
After the solar cell array and the antenna are unfolded, the dynamics of the satellite comprises the rotation of the star body and the elastic vibration and rotation of the solar cell array, and the star body attitude dynamics during attitude control of the thruster is as follows:
Figure BDA0002395263420000052
Figure BDA0002395263420000053
Figure BDA0002395263420000054
wherein, IsThe method comprises the following steps of (1) obtaining a rotational inertia matrix of a star in a three-axis coordinate system;
Figure BDA0002395263420000055
is the angular velocity vector of the star relative to the initial inertial coordinate system;
Figure BDA0002395263420000056
the angular velocity vectors of the left battery array and the right battery array are respectively; t issExternal moment acting on the satellite; fsls、FsrsFlexible coupling coefficients of vibration of the left battery array and the right battery array to rotation of the star body are respectively set; fals、FarsFlexible coupling coefficients of the vibration of the left battery array and the right battery array to the rotation of the battery arrays are respectively set; zetals、ζrsRespectively a left battery array modal damping coefficient and a right battery array modal damping coefficient; lambdaals、ΛarsRespectively a left battery array modal frequency matrix and a right battery array modal frequency matrix;
Figure BDA0002395263420000061
respectively a left solar cell array modal coordinate array and a right solar cell array modal coordinate array.
The xPC real-time operation module is in a real-time operation system environment, and outputs single-machine model data required by the system to the FPGA module by inputting peripheral dynamics module data and an operation instruction of the upper computer operation module. The xPC target machine is only used as an operating environment for model input and completes instruction operation.
Standalone models include, but are not limited to: attitude sensors such as star sensors and gyro combinations, and actuating components such as reaction flywheels.
Wherein, each single model can be expressed by a formula as:
(1) the star sensor model mainly considers the attitude calculation into quaternion output according to the attitude angle of an inertia system measured by the sensor, and simultaneously considers the influence of noise.
Model parameters are configured preliminarily according to experience, and then redesigning can be carried out according to specific requirements.
By inputting the attitude phi of the spacecraft in inertial spacei、θi、ψiAnd solving to obtain the attitude measurement value of the spacecraft in the inertial space (the direction Z of the optical axis in the inertial space)I)。
The optical axis is expressed as
Figure BDA0002395263420000062
Transformation matrix from body coordinate system to measurement coordinate system, CbsIs a constant matrix.
The optical axis is represented as Z under the systemb＝CbsZs；
Conversion matrix C from main system to inertial systemib；
The optical axis is directed in inertial space to ZI＝CIbZbFrom the inertial space attitude (phi) of the spacecrafti、θi、ψi) Calculated according to the main shaft rotation rule.
(2) The gyro combination model is mainly used for carrying out attitude calculation according to the measured attitude angular velocity under the inertial system, considering disturbance factors and amplitude limiting effects and calculating the attitude angular velocity after the factors are considered.
The amplitude limiting equation:
Figure BDA0002395263420000063
in the formula: omegaoutTo output the attitude angular velocity, rad/s; omegainTo input attitude angular velocity, rad/s; omegaminIs the minimum amplitude limit, rad/s; omegamaxFor maximum amplitude limitation, rad/s.
Noise calculation formula:
distr＝Aisin(2·π·w·t)
in the formula: distr is the disturbance magnitude, rad; a is the perturbation amplitude, rad; w is the perturbation frequency, rad/s; t is time, s.
(3) Reaction flywheel
Converting the angular momentum change rate of the reaction flywheel to three main inertia axes of the satellite through an installation matrix; the controller calculates the generated three-axis control torque, and after force pressure conversion, the distribution matrix D corresponding to the flywheel mounting matrix converts the three-axis torque instruction to each flywheel;
according to the optimal control principle, the distribution matrix is the pseudo-inverse of the installation matrix:
D＝RT(RRT)-1
in the formula: rTIs the transposed matrix of R.
The software of the simulation machine system can be mainly divided into an xPC real-time operating system module, an FPGA module, an upper computer control module, a dynamics module and the like.
a) xPC target machine module development
The Target machine starting kernel is used for calling and running the real-time operating system environment of the xPC Target, and the Target starting kernel needs to be regenerated or updated after the environment attribute of the xPC Target is changed every time. Under an xPC real-time simulation platform, different simulation systems correspond to different peripheral equipment, so that the technology for autonomously developing the I/O driving module is provided. For an xPC target, peripheral hardware is divided into two types, one type is xPC-supported equipment, and the xPC target can be used only by installing a corresponding board card drive; for the device unsupported by xPC, a user needs to utilize the C MEX S function to develop a driver program for a specific hardware device.
And (3) xPC application program development:
1. establishing a Simulink model;
2. adding an xPC Target Scope module (for realizing data visualization):
1) opening a Simulink Library Browser dialog box, and positioning to xPC Target → Misc;
2) and adding a scope (xPC) module for the model.
3) Parameters of the scope (xPC) module are set.
3. Adding an I/O device driving module:
1) and selecting the driving module according to the manufacturer and the model of the I/O equipment.
2) Setting drive module parameters
4. Setting simulation parameters: and selecting the Fix-step, and selecting the step length according to the actual situation.
5. Starting the target machine: the target machine is started in DOSLoader mode.
6. Setting RTW parameters:
1) open the RTW tab and select xpctarget.
2) The content in the xPC Target options selection is set (generally by default).
7. Creating and downloading target application
1) Click the Real-Time Workshop menu item under the Tools menu under the Simulink window, and click the BuildModel command. At which point the program creation process is performed.
2) Enter in the MATLAB command window: tg (target object name), view target object properties.
8. Running xPC Target program
After the xPC Target application program creating process and the downloading process are completed, an xPC Target object is generated, the object represents the Target computer and the Target application program, the xPC Target object is defined by a group of attributes and related methods, and the object Target attributes can be changed through an object Target method, so that the operation of the Target application program and the Target computer are controlled. The user can control the target program through four methods as follows:
1)xPC Target Explorer
2) MATLAB Command line
3) Simulink external mode (convenient parameter adjustment)
4) Web browser
9. Start is typed in the MATLAB command line to start the target program.
Stop can stop the target program. Parameters are directly controlled and modified in xPC Target Explorer. In Simulink by external mode control.
10. Generating independently running target applications
After the model is debugged, the target application program which runs independently can be generated through the StandAlone mode, so that the target application program can run independently without a host machine, and the physical simulation is really realized. The specific steps are that the xPCarget environment attribute TargetBOOT is set as StandAlone, then the Build Model is used, and a subdirectory is generated under the current directory, wherein the subdirectory comprises files such as Model. And copying the files to a starting disc, putting the starting disc into the target machine, and starting the target machine to run the target application program.
b) FPGA module design and implementation
The FPGA is a core processing device of the interface board card, and the control and simulation of the star sensor are realized through a plurality of serial ports, selection, insertion and other modules. The FPGA module is designed by adopting Verilog language. The design rules follow the top-down design methodology common to digital circuit design. The top module is a system related functional module, a hierarchical design is adopted downwards, and module division follows three principles of function classification, interface simplification and moderate scale.
The modules of the whole FPGA are decomposed into a plurality of independent modules, as shown in fig. 2, which are a real single machine interface module, a signal source selection module, a fault insertion module, an AOCC interface module, and a PCIe interface module, respectively.
Determining main function modules of FPGA software, and generating a top-layer module relationship as follows: sim _ sys _ top is the top layer of FPGA software, and the module is the top layer of FPGA software codes; realsys _ intf is a clock and reset module, which is used for generating a global clock and reset; realsys _ intf is a real single-machine interface module, is connected with a digital interface of a real single machine and can be obtained by counting the number of each functional single-machine interface of a task; source _ sel is a data source selection module, decomposes data of a real single machine interface, and can select real single machine data or target machine simulation data according to fields; err _ ins is a fault insertion module, and is used for generating and inserting (replacing) fault data for data ready to be reported to AOCC (automated optical storage controller), so that the purpose of simulating single machine fault is achieved. and AOCC _ intf is an AOCC interface and is responsible for reporting AOCC interface data. The pci _ intf is a PCIe interface module and is responsible for analyzing PCIe instructions and reporting state data.
c) Human-machine interface module development of control software
In order to conveniently call the data format of the MATLAB, the interface design of the control module uses a GUI design kit carried by the MATLAB, and the interface of the control software must support the functions of parameter display, selection of a real stand-alone model and a stand-alone model, fault insertion and the like, so that fig. 3 gives a schematic diagram of the control software, and the basic functions of the control software are divided into 4 parts.
1) The satellite state. The part mainly analyzes the received dynamic data and displays corresponding parameters in real time.
2) And (6) debugging and selecting. All real single machines or all single machine models can be selected through intuitive one-key operation, and corresponding overall debugging is carried out.
3) And (4) fault insertion. And independently selecting a real single machine or a single machine model, inserting and simulating fault parameters of a specific single machine, and setting a corresponding fault model by one key.
4) And (4) access state indication. If all real single machines (green) or all single machine models (yellow) are selected in the debugging selection, the single machines of the part display the colors corresponding to the real single machines or the single machine models. If a real single machine or a single machine model is independently selected in the fault insertion link, after each pair of independent single machines carry out parameter setting and send fault simulation, the color of the current single machine type is updated to be green or yellow in real time, and the monitoring requirement is met.
The embodiment also provides a full-function modular simulation processing method based on satellite attitude and orbit control, which comprises the following steps: the dynamic module transmits the single-machine model data to the xPC real-time operation module, and the upper computer control module transmits an operation instruction to the xPC real-time operation module; the xPC real-time operation module receives the single-machine model data and the operation instruction and transmits the single-machine model data to the FPGA module according to the operation instruction; the FPGA module obtains an inertial space attitude measurement value of the spacecraft, an output attitude angular velocity, noise and an installation matrix of a reaction flywheel according to the model data of the single machine, and transmits the installation matrix to the real satellite-borne single machine.
Compared with real-time operating systems such as dSPACE and LabVIEW, the invention has richer I/O resources and flexibly writes the required driving module by the characteristics of the xPC real-time operating system, and has low hardware cost and high cost performance.
The invention adopts the unified design through all the interface boards, comprehensively considers the characteristics of different single machines, selects the proper main control chip and realizes the information processing and switching of different single machines by loading different programs. The effect that reaches does: the hardware design is unified, the hardware design of each interface board card is unified, the hardware design of the board card is simplified, the reliability of an electric appliance of the board card can be effectively improved, and the risk is reduced; the system has strong flexibility and the universal design ensures that different single-computer functions are realized and modified through software. This provides greater openness to future changes in the number or functionality of the individual machines.
The invention carries out software design and realization through the FPGA, and the FPGA has incomparable advantages in various logic devices such as the FPGA, a DSP, an MCU and the like due to the large demand of the simulator system on I/O. Besides the advantages of rich interfaces and good sequential control, the I/O quantity of the FPGA has great advantages in magnitude order compared with DSP, MCU and the like. The timing precision and controllability of the FPGA are far higher than those of other devices. Therefore, the FPGA is selected as a main control chip, and a plurality of completely consistent interface board cards are uniformly designed to be the best scheme of the simulation machine system.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make variations and modifications of the present invention without departing from the spirit and scope of the present invention by using the methods and technical contents disclosed above.

Claims (10)

1. A full-function modularization simulation processing system based on satellite attitude and orbit control is characterized by comprising: the system comprises a dynamics module, an upper computer control module, an xPC real-time operation module and an FPGA module; wherein the content of the first and second substances,
the dynamic module transmits single-machine model data to the xPC real-time operation module, and the upper computer control module transmits an operation instruction to the xPC real-time operation module;
the xPC real-time operation module receives the single-machine model data and the operation instruction and transmits the single-machine model data to the FPGA module according to the operation instruction;
and the FPGA module obtains an inertial space attitude measurement value of the spacecraft, an output attitude angular velocity, noise and an installation matrix of a reaction flywheel according to the single-machine model data and transmits the installation matrix to the real satellite-borne single machine.
2. The fully functional modular satellite attitude and orbit control-based analog processing system of claim 1, further comprising: the data of the single-unit model is obtained through a solar cell array and a star attitude dynamic equation during attitude control of the thruster after the antenna is unfolded.
3. The fully functional modular satellite attitude and orbit control-based analog processing system of claim 2, further comprising: the dynamic equation of the star attitude during attitude control of the thruster after the solar cell array and the antenna are unfolded is as follows:
Figure FDA0002395263410000011
Figure FDA0002395263410000012
Figure FDA0002395263410000013
wherein, IsThe method comprises the following steps of (1) obtaining a rotational inertia matrix of a star in a three-axis coordinate system;
Figure FDA0002395263410000014
is the angular velocity vector of the star relative to the initial inertial coordinate system;
Figure FDA0002395263410000015
the angular velocity vectors of the left battery array and the right battery array are respectively; t issExternal moment acting on the satellite; fsls、FsrsFlexible coupling coefficients of vibration of the left battery array and the right battery array to rotation of the star body are respectively set; fals、FarsFlexible coupling coefficients of the vibration of the left battery array and the right battery array to the rotation of the battery arrays are respectively set; zetals、ζrsRespectively a left battery array modal damping coefficient and a right battery array modal damping coefficient; lambdaals、ΛarsRespectively a left battery array modal frequency matrix and a right battery array modal frequency matrix;
Figure FDA0002395263410000016
respectively a left solar cell array modal coordinate array and a right solar cell array modal coordinate array.
4. The fully functional modular satellite attitude and orbit control-based analog processing system of claim 1, further comprising: the attitude measurement value of the spacecraft in the inertial space is the pointing direction Z of an optical axis in the inertial spaceI＝CIbZbWherein, CibBeing body-tied to inertial systemTransformation matrix, ZbIs represented by an optical axis under the system, Zb＝CbsZs，CbsIs a matrix of constant values, and the matrix of constant values,
Figure FDA0002395263410000021
Zsthe optical axis is expressed in the star sensor measurement coordinate system.
5. The fully functional modular satellite attitude and orbit control-based analog processing system of claim 1, further comprising: the output attitude angular velocity is:
Figure FDA0002395263410000022
wherein, ω isoutTo output attitude angular velocity, omegainFor inputting attitude angular velocity, omegaminTo a minimum amplitude limit, ωmaxIs the maximum amplitude limit.
6. The fully functional modular satellite attitude and orbit control-based analog processing system of claim 1, further comprising: the formula for noise is:
distr＝A·sin(2·π·w·t)；
where distr is noise, A is a disturbance amplitude, w is a disturbance frequency, and t is time.
7. The fully functional modular satellite attitude and orbit control-based analog processing system of claim 1, further comprising: the mounting matrix of the reaction flywheel is obtained by the following formula:
D＝RT(RRT)-1；
wherein D is the distribution matrix and R is the mounting matrix of the reaction flywheel.
8. A full-function modular simulation processing method based on satellite attitude and orbit control is characterized by comprising the following steps:
the dynamic module transmits the single-machine model data to the xPC real-time operation module, and the upper computer control module transmits an operation instruction to the xPC real-time operation module;
the xPC real-time operation module receives the single-machine model data and the operation instruction and transmits the single-machine model data to the FPGA module according to the operation instruction;
the FPGA module obtains an inertial space attitude measurement value of the spacecraft, an output attitude angular velocity, noise and an installation matrix of a reaction flywheel according to the model data of the single machine, and transmits the installation matrix to the real satellite-borne single machine.
9. The full-function modular simulation processing method based on satellite attitude and orbit control of claim 8, characterized in that: the data of the single-unit model is obtained through a solar cell array and a star attitude dynamic equation during attitude control of the thruster after the antenna is unfolded.
10. The full-function modular simulation processing method based on satellite attitude and orbit control of claim 8, characterized in that: the dynamic equation of the star attitude during attitude control of the thruster after the solar cell array and the antenna are unfolded is as follows:
Figure FDA0002395263410000031
Figure FDA0002395263410000032
Figure FDA0002395263410000033
wherein, IsThe method comprises the following steps of (1) obtaining a rotational inertia matrix of a star in a three-axis coordinate system;
Figure FDA0002395263410000034
is the angular velocity vector of the star relative to the initial inertial coordinate system;
Figure FDA0002395263410000035
the angular velocity vectors of the left battery array and the right battery array are respectively; t issExternal moment acting on the satellite; fsls、FsrsFlexible coupling coefficients of vibration of the left battery array and the right battery array to rotation of the star body are respectively set; fals、FarsFlexible coupling coefficients of the vibration of the left battery array and the right battery array to the rotation of the battery arrays are respectively set; zetals、ζrsRespectively a left battery array modal damping coefficient and a right battery array modal damping coefficient; lambdaals、ΛarsRespectively a left battery array modal frequency matrix and a right battery array modal frequency matrix;
Figure FDA0002395263410000036
respectively a left solar cell array modal coordinate array and a right solar cell array modal coordinate array.
CN202010128913.7A 2020-02-28 2020-02-28 Full-function modularized simulation processing system and method based on satellite attitude and orbit control Active CN111308911B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202010128913.7A CN111308911B (en) 2020-02-28 2020-02-28 Full-function modularized simulation processing system and method based on satellite attitude and orbit control

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202010128913.7A CN111308911B (en) 2020-02-28 2020-02-28 Full-function modularized simulation processing system and method based on satellite attitude and orbit control

Publications (2)

Publication Number Publication Date
CN111308911A true CN111308911A (en) 2020-06-19
CN111308911B CN111308911B (en) 2023-05-09

Family

ID=71159482

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202010128913.7A Active CN111308911B (en) 2020-02-28 2020-02-28 Full-function modularized simulation processing system and method based on satellite attitude and orbit control

Country Status (1)

Country Link
CN (1) CN111308911B (en)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112034731A (en) * 2020-08-12 2020-12-04 中国科学院国家空间科学中心 Payload semi-physical simulation system based on associated knowledge
CN112596504A (en) * 2020-12-22 2021-04-02 上海卫星工程研究所 Flywheel signal acquisition equipment for satellite attitude and orbit control general comprehensive test
CN112666847A (en) * 2021-03-18 2021-04-16 长沙天仪空间科技研究院有限公司 Satellite test system and method suitable for various simulation states
CN112947521A (en) * 2021-02-10 2021-06-11 西南电子技术研究所（中国电子科技集团公司第十研究所） Multifunctional simulation platform of spacecraft measurement and control system
CN114212279A (en) * 2021-11-30 2022-03-22 上海航天控制技术研究所 Angular momentum management method for redundantly installed flywheel

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20070129922A1 (en) * 2005-12-01 2007-06-07 Electronics And Telecommunications Research Institute Satellite simulation system using component-based satellite modeling
CN102495552A (en) * 2011-12-08 2012-06-13 哈尔滨工业大学 Real-time simulation system oriented to space-borne electronic system
CN103092212A (en) * 2013-01-08 2013-05-08 天津大学 Emulation verification platform and implementing method of micro-satellite team formation system
CN105974907A (en) * 2016-05-12 2016-09-28 上海微小卫星工程中心 Satellite attitude control ground simulation testing system
CN109522622A (en) * 2018-10-31 2019-03-26 中国运载火箭技术研究院 A kind of in-orbit load working condition of multiple degrees of freedom solar battery array determines method and system
US20190302259A1 (en) * 2018-03-27 2019-10-03 The Mathworks, Inc. Systems and methods for generating synthetic sensor data
CN110442043A (en) * 2019-08-14 2019-11-12 上海卫星工程研究所 Satellite GNC system in real time-semi physical flexibility emulation test system
CN110471311A (en) * 2019-08-15 2019-11-19 南京理工大学 A kind of micro-nano satellite rail control semi-physical simulation synthesis experiment platform

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20070129922A1 (en) * 2005-12-01 2007-06-07 Electronics And Telecommunications Research Institute Satellite simulation system using component-based satellite modeling
CN102495552A (en) * 2011-12-08 2012-06-13 哈尔滨工业大学 Real-time simulation system oriented to space-borne electronic system
CN103092212A (en) * 2013-01-08 2013-05-08 天津大学 Emulation verification platform and implementing method of micro-satellite team formation system
CN105974907A (en) * 2016-05-12 2016-09-28 上海微小卫星工程中心 Satellite attitude control ground simulation testing system
US20190302259A1 (en) * 2018-03-27 2019-10-03 The Mathworks, Inc. Systems and methods for generating synthetic sensor data
CN109522622A (en) * 2018-10-31 2019-03-26 中国运载火箭技术研究院 A kind of in-orbit load working condition of multiple degrees of freedom solar battery array determines method and system
CN110442043A (en) * 2019-08-14 2019-11-12 上海卫星工程研究所 Satellite GNC system in real time-semi physical flexibility emulation test system
CN110471311A (en) * 2019-08-15 2019-11-19 南京理工大学 A kind of micro-nano satellite rail control semi-physical simulation synthesis experiment platform

Non-Patent Citations (11)

* Cited by examiner, † Cited by third party
Title
JIAN SUN: "Research on Physical Simulation Experiments for Satellite Multi-axis Pointing Control", 《PROCEEDINGS OF THE 2007 IEEE INTERNATIONAL CONFERENCE ON ROBOTICS AND BIOMIMETICS》 *
REURISON RODRIGUES: "Hardware in the Loop Simulation for Model Predictive Control Applied to Satellite Attitude Control", 《IEEE ACCESS》 *
张小伟: "用于长周期高精度轨道控制任务的快速半实物仿真系统", 《空间控制技术与应用》 *
张庆君: "《国之重器出版工程 航天器系统设计》", 31 May 2018 *
张怡文: "卫星姿轨控半物理仿真测试系统", 《计算机测量与控制》 *
杜耀珂: "基于ＳＴＫ的卫星实时视景仿真系统设计", 《空间控制技术与应用》 *
杨正贤;孔宪仁;王继河;邱文勋;王峰;: "基于xPC的小卫星半物理仿真验证平台" *
王嘉轶: "Fuzzy PID Controller based on Improved Neural Network for Satellite Attitude", 《MEASUREMENT, COMPUTER, COMMUNICATION》 *
胡军: "《特征建模理论、方法和应用》", 31 October 2019 *
胡敏;吴锦杰;曾国强;: "基于自组织网络的航天器姿态控制半实物仿真" *
解永春: "《航天器动力学与控制》", 31 March 2018 *

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112034731A (en) * 2020-08-12 2020-12-04 中国科学院国家空间科学中心 Payload semi-physical simulation system based on associated knowledge
CN112596504A (en) * 2020-12-22 2021-04-02 上海卫星工程研究所 Flywheel signal acquisition equipment for satellite attitude and orbit control general comprehensive test
CN112947521A (en) * 2021-02-10 2021-06-11 西南电子技术研究所（中国电子科技集团公司第十研究所） Multifunctional simulation platform of spacecraft measurement and control system
CN112666847A (en) * 2021-03-18 2021-04-16 长沙天仪空间科技研究院有限公司 Satellite test system and method suitable for various simulation states
CN114212279A (en) * 2021-11-30 2022-03-22 上海航天控制技术研究所 Angular momentum management method for redundantly installed flywheel
CN114212279B (en) * 2021-11-30 2023-11-10 上海航天控制技术研究所 Angular momentum management method for redundant installation of flywheel

Also Published As

Publication number Publication date
CN111308911B (en) 2023-05-09

Similar Documents

Publication Publication Date Title
CN111308911B (en) Full-function modularized simulation processing system and method based on satellite attitude and orbit control
Maclay Simulation gets into the loop
CN108228467B (en) Embedded flight control software algorithm rapid verification method and system
US10552560B2 (en) Controlling real time during embedded system development
JP2008170998A (en) System and method for turbine control simulation
CN105259794B (en) A kind of satellite gravity anomaly ground artificial system based on plug and play technique
KR101862221B1 (en) Flight control law simulation method and apparatus
CN102436184B (en) EMP (embedded microprocessor)-based double-target machine control real-time simulation system
CN112214902B (en) Real-time simulation system for attitude and orbit control and single-machine communication of satellite
CN111638654A (en) Fault-adaptive intelligent control semi-physical simulation method for carrier rocket
CN112947384A (en) Multifunctional satellite simulation test system and method
CN110383186A (en) A kind of emulation mode and device of unmanned plane
CN112131741B (en) Real-time dual-kernel single-machine semi-physical simulation architecture and simulation method
CN112099518A (en) Satellite attitude control simulation test method and system
Zhu et al. A new hierarchical software architecture towards safety-critical aspects of a drone system
Gholkar et al. Hardware-in-loop simulator for mini aerial vehicle
Sanvido Hardware-in-the-loop Simulation Framework
CN112632700B (en) Flywheel signal acquisition method and system for satellite attitude and orbit control general comprehensive test
Horowitz et al. Embedded software design and system integration for rotorcraft UAV using platforms
CN109471424B (en) Ground simulation test system and method for large-angle momentum rotating part
Shixianjun et al. Hardware-in-the-loop simulation framework design for a UAV embedded control system
Rolfsmeier et al. A new calibration system for ECU development
Reinholtz Applying simulation to the development of spacecraft flight software
Wang et al. Design of Hardware-in-the-loop Simulation System based on RTX and FlightSim™
CN112550771B (en) Control moment gyro signal acquisition method and system for attitude and orbit control general comprehensive test

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant